Aircraft using turbo-electric hybrid propulsion system

ABSTRACT

An air vehicle incorporating a hybrid propulsion system. The system includes a gas turbine engine as a first motive power source, and one or more battery packs as a second motive power source. Through selective coupling to a DC electric motor that can in turn be connected to a bladed rotor or other lift-producing device, the motive sources provide differing ways in which an aircraft can operate. In one example, the gas turbine engine can provide operation for a majority of the flight envelope of the aircraft, while the battery packs can provide operation during such times when gas turbine-based motive power is unavailable or particularly disadvantageous. In another example, both sources of motive power may be decoupled from the bladed rotor such that the vehicle can operate as an autogyro.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a divisional of pending (and now-allowed) U.S.patent application Ser. No. 11/972,879, filed Jan. 11, 2008.

BACKGROUND OF THE INVENTION

The present invention relates generally to a hybrid propulsion systemfor aircraft, and more particularly to a gas turbine-based propulsionsystem that can provide primary motive power to the aircraft over aportion of the aircraft's flight envelope, and a stored electric-basedpower system that is selectively coupled to the gas turbine such thatduring other portions of the aircraft's flight envelope, theelectric-based power system can provide the aircraft with its primarymotive power.

Aircraft are broadly categorized as either fixed-wing vehicles (such asan airplane) or rotary-wing vehicles (such helicopters and autogyros,the latter also referred to as autogiros or gyrocopters). Gas turbineengines are widely used to power both the fixed-wing and rotary wingforms of aircraft, where fixed-wing vehicles often employ turbofan,turbojet and turboprop variants, and rotary-wing vehicles often employturboshaft variants. In all circumstances, the basic gas generatorhardware is common, including a compressor, a combustor and turbine,where the compressor and turbine rotate on a generally common shaft (orset of concentric shafts) such that energy extracted from the turbine isused to power the compressor. Turbofans are very similar to turbojets,with the exception that they typically include an additional fan locatedupstream of the compressor. A turboprop engine, in addition to includingthe respective turbofan or turbojet componentry, also includes afore-mounted drive shaft that spins in common with the shaft of thecompressor and turbine. To match the high rotational speed of thecompressor and turbine to that of a propeller, a gearbox is insertedbetween the front end of the drive shaft and a propeller shaft.Turboshafts also include similar components to the turbofans andturbojets, and additionally include a shaft rotatably responsive toanother turbine stage.

In turboshaft engines, power generated by the gas generator (which isspinning about a generally horizontal axis) is transferred to the shaft(which is spinning about a generally vertical axis such that it can turna rotor made up of a series of blades that radially extend from acentral hub) through a gearing mechanism, such as a bevelled or wormgear. Shaft horsepower needs to produce a particular rotor rotationalspeed varies depending on the aircraft type, size and intended mission.For example, the CH-47 Chinook is a popular twin-rotor helicopterdesigned for commercial and military heavy lifting. Rotor blade powerrequirements for helicopters such as this may be in the range of fivethousand horsepower, while speed requirements of around two hundred andtwenty five revolutions per minute (RPM) are typical.

Despite their widespread use, conventional gas turbine-based propulsionsystems have significant drawbacks for certain types of aircraft. Forexample, in the event a turboshaft engine fails, a helicopter,gyrocopter or other inherently unstable aircraft has no way of returningto earth under its own power, and at best can expect to have to endure acontrolled crash landing. Likewise, if a helicopter employing aturboshaft engine as propulsive power is flying or hovering over an areawhere terrorists, armed conflict or related hostilities exist, theextreme heat put out by the engine or engines may make the helicopterexceedingly vulnerable to attack from infrared (IR) seeking weaponry.Accordingly, there exists a need for a propulsion system that overcomesthese shortcomings.

BRIEF SUMMARY OF THE INVENTION

These needs are met by the present invention, where in accordance with afirst aspect of the present invention, an aircraft employing a hybridpropulsion system is disclosed. The aircraft includes a fuselage, bodyor related airframe, a thrust-producing device coupled to the fuselage,and a hybrid propulsion system configured to provide power to operatethe thrust producing device. Cooperation between the propulsion systemand the thrust producing device provides motive power to the aircraft.In one form, the thrust-producing device is a bladed rotor, where therotating blades produce both lift and thrust if oriented properlyrelative to the aircraft. In the present context, the terms “lift” and“thrust”, while recognized in general aeronautical terms as representingtwo of the four primary forces acting upon an aircraft in flight (theother two being drag and weight), are used somewhat interchangeably asthose forces that contribute to the craft's upward or forward movement.Circumstances in the present disclosure where lift and thrust retaintheir traditionally-accepted aeronautical definitions will be apparentfrom the context. For example, the rotation of a bladed rotor that iscoupled to a helicopter will be understood to provide one or both oflift and thrust, depending at least in part on the orientation of therotor relative to the aircraft to which it is attached.

The propulsion system attains its hybrid nature by possessing two formsof power. The first comes from a gas turbine (i.e., jet) engine, whilethe second comes from an electric storage device. Such hybrid powersources may be tailored for use over various parts of the aircraftflight envelope where each exhibits relative strengths or advantages. Inthe present context, the flight envelope includes various parts of theflight path that the aircraft may be expected to encounter over thecourse of its operation. By way of non-limiting example, such partsinclude startup, ground loiter, takeoff, cruise, loiter/hover andlanding. As such, the gas turbine engine can be used to provide power tothe thrust and/or lift producing device over a portion of the aircraft'sflight envelope that requires long-term power, such as during aircraftcruise. Likewise, the electric storage device can be used to power theaircraft over a portion of the aircraft's flight envelope that benefitsfrom reduced thermal or pollutant emissions from the gas turbine engine,such as hovering, loitering or the like. An electrical generator is alsoplaced in cooperation with the shaft so that mechanical power from thegas turbine is converted to electric power for use by an electric motorthat is coupled to the thrust producing device. The presence of theelectrical generator allows the electric motor to accept a common formof power (i.e., electric current) from either the gas turbine engine orthe electric storage device.

Optionally, the electric storage device comprises at least one battery,which may be in the form of a single battery or part of a battery pack.The gas turbine engine may further include a transmission shaft thatextends from its main rotational shaft (i.e., the one or ones coupled tothe engine's compressor or turbine) to deliver power to the electricalgenerator. A clutch can be disposed on or with the shaft such thatduring the portion of the aircraft's flight envelope where that powergenerated in the gas turbine engine is to be used, the clutch providesthe necessary connectivity between the engine and the electricalgenerator. The battery or other electric storage can be kept chargedduring the portion of the aircraft's flight envelope where the gasturbine engine is providing the aircraft's power, as excess power fromthe gas turbine engine can be fed through the electrical generator tothe battery. In one form, the electric motor is a DC motor. As statedabove, the thrust producing device can be a bladed rotor oriented insuch a way that upon attaining a minimum rotational speed, it possessesat least one of lift and thrust attributes. During the portion of theaircraft's flight envelope where the power to operate the thrustproducing device is coming from the electric storage device, the gasturbine engine can be rendered substantially inoperable such that itsoperation is reduced or turned off entirely, where in the presentcontext, the term “substantially” refers to an arrangement of elementsor features that, while in theory would be expected to exhibit exactcorrespondence or behavior, may, in practice embody something less thanexact. As such, the term denotes the degree by which a quantitativevalue, measurement or other related representation may vary from astated reference without resulting in a change in the basic function ofthe subject matter at issue. A controller may be used to provide someautomation to the operation of the aircraft. Included is the ability tovary operation of the aircraft between a first and second portion of itsflight envelope. In a more particular form, the controller can be usedin conjunction with actuation equipment to provide full robotic controlof the aircraft over many or all segments of the aircraft flightenvelope, including take-off and landing. Such full control isespecially beneficial in unmanned aircraft configurations.

According to another aspect of the present invention, a rotary wingaircraft is disclosed. The aircraft includes a fuselage, one or morebladed rotors cooperative with the fuselage, and a hybrid propulsionsystem coupled to the fuselage and the one or more bladed rotors. Thepropulsion system includes a gas turbine engine, an electricalgenerator, an electric storage device and an electric motor selectivelycoupled to at least one of the electrical generator and the electricstorage device. The generator is coupled to the gas turbine engine sothat mechanical power received from the engine can be converted by theelectrical generator to electric power for use by the electric motor. Inthis way, the generator normalizes the form of energy being input intothe electric motor. During one (for example, a first) portion of aflight envelope of the aircraft, the gas turbine engine and the one ormore rotors provide at least a majority of motive power to the aircraft,while during another (for example, the second) portion of the flightenvelope, the electric storage device and the rotor or rotors provide atleast a majority of the motive power to the aircraft.

Optionally, the aircraft is a helicopter or gyrocopter. In eitherconfiguration, the aircraft may include numerous lift rotors. Suchrotors may be in tandem (i.e., front to back) or side-by-side, as wellas stacked. As before, the electric storage device may be made up of oneor more batteries. As with the previous aspect of the invention, therotary wing aircraft may include fully automated control throughoperation of a controller (for example, a computer-based device) withassociated actuation equipment to robotically manage some or allportions of the aircraft mission. Such controller and ancillaryequipment could also be configured to cooperate with a transceiver orrelated communication system to allow for remote control of theaircraft.

According to another aspect of the present invention, a method ofproviding motive power to an aircraft is disclosed. The method includescoupling a hybrid propulsion system to a thrust producing device andoperating the system such that power produced by it operates the thrustproducing device. The hybrid propulsion system includes a gas turbineengine, an electric storage device configured to deliver an electriccurrent and an electrical generator responsive to the gas turbine enginesuch that power produced by the operation of the engine is convertedinto electric current. In addition, the system includes an electricmotor selectively responsive to the generator and the electric storagedevice. By having the electric motor be responsive to electric currentirrespective of its source (i.e., the gas turbine engine or the electricstorage device), the use of redundant componentry is avoided.

Optionally, the thrust producing device operates by rotating in responseto power provided to it by the electric motor. Furthermore, the thrustproducing device is configured to provide both lift and thrust to theaircraft, which may be configured as a helicopter or gyrocopter. Inaddition, the electric storage device can be made up of one or morebatteries. In a particular form, the operation of one of the gas turbineengine and the electric storage device comprises operating the electricstorage device substantially exclusively. During such mode of operation,the gas turbine engine can be turned off. In a related mode ofoperation, operation of the gas turbine engine can be curtailed enoughso that one or more of its thermal output, pollutant output and abilityto produce useful thrust for the aircraft is substantially ceased. Inaddition, the operation of one of the gas turbine engine and at leastone battery comprises operating the gas turbine engine such that excesspower produced by it is conveyed to the electric storage device throughthe electrical generator to maintain the electric storage device in asubstantially charged condition. As with the previous aspects,computerized control may be affected through a controller and associatedactuators, as well as through cooperation of the same with a transceiverto allow remote control, depending on the mission, aircraftconfiguration or like considerations.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The following detailed description of the present invention can be bestunderstood when read in conjunction with the following drawings, wherelike structure is indicated with like reference numerals and in which:

FIG. 1 shows schematically a hybrid propulsion system according to anembodiment of the present invention;

FIG. 2A shows a top view of a rotary-winged aircraft with a single mainrotor and horizontal stability rotor, utilizing a pair of the hybridpropulsion systems of FIG. 1;

FIG. 2B shows a front elevation view of the rotary-winged aircraft ofFIG. 2A;

FIG. 3A shows a top view of a rotary-winged aircraft with afront-to-back dual rotor configuration, utilizing four of the hybridpropulsion systems of FIG. 1;

FIG. 3B shows a front elevation view of the rotary-winged aircraft ofFIG. 3A;

FIG. 4A shows a top view of a rotary-winged aircraft with a side-to-sidedual rotor configuration, utilizing four of the hybrid propulsionsystems of FIG. 1;

FIG. 4B shows a front elevation view of the rotary-winged aircraft ofFIG. 4A;

FIG. 5A shows a top view of a rotary-winged aircraft with a side-to-sideand front-to-back quadruple rotor configuration, utilizing eight of thehybrid propulsion systems of FIG. 1;

FIG. 5B shows a front elevation view of the rotary-winged aircraft ofFIG. 5A;

FIG. 6A shows a top view of a rotary-winged aircraft with a single rotorauto gyro configuration utilizing a pair of the hybrid propulsionsystems of FIG. 1;

FIG. 6B shows a front elevation view of the rotary-winged aircraft ofFIG. 6A;

FIG. 7 shows schematically a hybrid propulsion system according to analternate embodiment of the present invention;

FIG. 8A shows a top view of a rotary-winged aircraft with a side-to-sidedual rotor configuration, utilizing four of the hybrid propulsionsystems of FIG. 7; and

FIG. 8B shows a front elevation view of the rotary-winged aircraft ofFIG. 8A

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring initially to FIGS. 1 and 7, a hybrid propulsion system 1according to two embodiments of the present invention is shown. In thepresent context, a hybrid propulsion system is one that provides motivepower for an aircraft from two or more disparate power sources. Theinvention disclosed herein achieves that objective through shaft powerprovided by an internal combustion engine and electric power provided byone or more batteries. Both power sources, which are described below inmore detail, can convert their energy into a form useable by an electricmotor that can be used to turn one or more lifting surfaces on anaircraft. In a more particular form, the hybrid propulsion system 1 is aturbo-electric propulsion system that is particularly well-suited topowering a rotary-winged aircraft such as a helicopter or autogyro.Referring with particularity to FIG. 1, the components making up hybridpropulsion system 1 are arranged in a lateral configuration, whilereferring with particularity to FIG. 7, the components making up hybridpropulsion system 1 are arranged in an in-line configuration. The choiceof which configuration is dictated by weight, volume and relatedaircraft integration concerns, among other things. As such, either ofthe hybrid propulsion systems 7 depicted in FIGS. 1 and 7 are applicableto any of the aircraft shown in the remaining drawings, subject toconfigurational limitations imposed by the aircraft.

System 1 includes as the first of these motive power sources a gasturbine engine 5 with inlet 5A and exhaust 5B. An electric generator 10is connected to the engine 5 through a primary transmission shaft 15that extends perpendicularly between them so that rotational movement ofthe engine shaft 5C imparts comparable rotational movement within thegenerator 10. Generator 10 can be of either an alternating current (AC)or direct current (DC) variety, although the remainder of thisdisclosure will focus on the DC variant, since such a configuration willabrogate the need for a battery-charging rectifier. Primary transmissionshaft 15 is shown in simplified fashion as extending directly fromengine shaft 5C in front of engine inlet 5A for ease of visualization.As will be appreciated by those skilled in the art, primary transmissionshaft 15 is preferably in the form of a power take-off shaft that passesthrough the inside of an inlet guide vane (not shown) or strut of theengine's fan or compressor section (neither of which are shown). Aclutch 20 is used to allow the engine 5 and generator 10 to beselectively decoupled from one another at the primary transmission shaft15. In one form, clutch 20 utilizes a mechanical friction plate tofacilitate the necessary connection and disconnection. Suchdisconnection allows the engine 5 to operate independently for jetflight (which may be particularly useful for fixed wing aircraft).Connection of the primary transmission shaft 15 to DC electric generator10, as well as connection of the primary transmission shaft 15 to theengine 5 can be effected by known means, such as a worm gear connection,bevel gear connection or the like. The DC electric generator 10 iscapable of supplying adequate DC current to run the DC electric motor25, which in turn can supply sufficient power to turn a propeller (orrelated bladed rotor) 30 (presently shown with only blade roots) that isused to provide lift to the aircraft.

The second source of motive power may come from one or more externalbattery packs 35. Rather than converting the rotational mechanicalenergy of the engine 5 and primary transmission shaft 15 into DCelectric current through the generator 10 to provide the DC electricmotor 25, the external battery packs 35 provide it directly to the motor25. While such a battery pack 35 may be made large enough to provideaircraft motive power for relatively long durations of flight, thepresent inventor has recognized that a battery pack 35 so sized may beimpractically heavy, bulky or the like. As such, the battery pack 35should be sized to provide such motive power for short-term operation,the duration of which is dictated by anticipated mission or emergencyrequirements.

Switching between the two power sources of engine 5 and external batterypack 35 is effected through a controller 40 that may be connected to oneor more actuators (not shown) through either cables 45 for hardwiredconnection or a wireless transceiver 50. For example, if a determinationis made (such as through pilot input, for example) that motive power isto be provided to the aircraft through the external battery pack 35, thecontroller 40 can instruct the clutch 20 to decouple the primarytransmission shaft 15 from engine 5. Such disconnection may beparticularly advantageous in circumstances where operation of the engine5 is inoperable or potentially harmful or wasteful. For example,powering the bladed rotor 30 through electricity provided by theexternal battery pack 35 can be used to run an aircraft in a low thermaloutput mode until such time as it is determined appropriate to rely uponthe engine 5 for such motive power. In one form, controller 40 is acomputer (i.e., microprocessor)-based system that can control some orall aspects of the aircraft flight envelope. Such allows preprogrammedautomation of actuator functions, as well as the ability to accept inputcontrol signals from a remote location, in the case of unmanned airvehicles (UAVs), through transceiver 50.

A secondary transmission shaft 55 can be used to supply power toadditional rotor(s), such as a rotor for horizontal stabilization (inthe case of a single rotor helicopter or the like), or a second liftingrotor in the case of a tandem-rotor helicopter. Another clutch 60 can beused in a manner generally similar to clutch 20 to selectively couplethe DC electric motor 25 and bladed rotor 30. Disconnection of bladedrotor 30 from the engine 5 and battery pack 35 through clutch 60 permitsthe bladed rotor 30 to freely spin, which may be advantageous duringemergency procedures as it allows the aircraft to operate in an autogyromode with a more controlled descent. Likewise, clutch 60 can enable theaircraft to run on electric power supplied by an appropriately sizedinternal battery pack (discussed below) for start-up and other groundoperations.

As will be appreciated by those skilled in the art, operation of theengine 5 to turn the bladed rotor 30 involves the conversion ofmechanical (shaft) energy in the engine 5 to electrical energy in thegenerator 10 and back into mechanical energy at the motor 25. Reductionsin propulsion system 1 efficiency due to losses attendant to eachconversion are more than offset by the increased functionality thatarises out of having both the engine 5 and external battery pack 35provide motive power to the aircraft. The present inventor hasdiscovered that at least for UAVs, autogyros, small single rotorhelicopters and related craft, the present invention an related benefitsprovided thereby is a practical way to achieve desirable systemredundancy and multiple modes of operation.

An internal battery pack 65 capable of providing start up power andpowering any needed internal systems is also provided. Such internalbattery pack 65 may be connected to an auxiliary power unit (APU, notshown) to effect such starting and aircraft support functions. Inaddition to clutches 20 and 60, controller 40, which for example is amicroprocessor-based system, may among other things, control operationof the battery packs 35 and 65, speed of the DC electric motor 25, andpitch of bladed rotor 30. Wireless transceiver 50 may, in addition toproviding local controls within the aircraft, receive fly-by-wirecommands for unmanned operation, such as in a UAV. In such a case, thewireless transceiver 50 and controller 40 may cooperate to providecomplete automated electronic (i.e., computerized) robotic control ofthe aircraft over a part or the entirety of its mission. In a latterexample, such robotic control may include take-off and landing, as wellas in-flight maneuvers. Cable 45 or related wiring, such as thatfamiliar to those skilled in the art, is used to provide electricalconnectivity between the various components. Such connectivity can beused to provide low current control and information signals, as well ashigh current to the DC electric motor 25 and related motive powercomponentry.

There are at least three types of DC electric motors 25 available,generally categorized as traditional, brushless, and coreless.Traditional DC motors use a core of iron bound with copper or otherhighly conductive wire. The core is centered in a series of magnets suchthat power (in the form of electric current) is transmitted to the corevia graphite brushes. In the present context, the electric current is DCcoming from the generator 10 or battery pack 35. Brushless motors usemany of the same materials as the traditional DC motor configuration,but reverse them, placing the magnets in the center surrounded by thecopper wiring. Power is supplied to the copper wiring, which allows themagnet to spin. The coreless motors are arranged in the same manner astraditional DC motors, but replace the iron and copper with aluminummeshes bound in glass epoxy. The choice of motor 25 configuration can bereadily made depending on (among other factors) the type of application,as well as motor availability, weight and cost. These factors beingequal, the brushless type is preferred for both the generator and therotor motors due to the brushless motor's greater reliability andlongevity.

Traditional DC electric motors, with significant amounts of metallicwiring wrapped around magnetizable poles, tend to be heavy. To reducethe weight associated with such motors, the present inventor hasdiscovered that DC electric motors 25 could be tailored to specificparts of the envelope, where maximum power (and concomitant motor size)may not be required. For example, if the DC electric motors 25 could beused for limited periods in relatively lightly-loaded segments (such ashover in the case of a helicopter), then such could provide separateutility over the power coming from the gas turbine engine 5 that wouldbe used for take-off or related maneuvers that require maximum amountsof power. In a related way, the DC electric motor 25 can be configuredto not be the sole source of motive power. In such circumstances,smaller, lighter weight motors can be used effectively. Otherweight-reducing schemes could further be employed. For example, themajority of the weight of a DC electric motor 25 comes from the motorhousing, which is usually steel or a related iron-based metal. Suchweight could be significantly reduced by using a lightweight, compositematerial for the motor housing such as high heat resistant thermosetresins with fiber glass reinforcement or newer, high strength, high heatresistant organic film. Further weight reduction would be possible byreplacing the iron core magnets with rare earth cobalt (or related)magnets. Other options, such as nickel-metal hydride and lithium ionbatteries used to provide power to hybrid automobiles could be used toprovide the necessary power without the significant weight penalties oftraditional electric motors. The power requirements of the engines arebased on the size and configuration of the aircraft. For example, a U.S.Chinook helicopter requires about 4900 shaft horsepower to turn theblades at approximately 225 revolutions per minute, generating around113600 foot-pounds of torque. Likewise, a U.S. Black Hawk helicopterrequires about 1900 shaft horsepower to turn the blades at approximately260 revolutions per minute, generating around 39500 foot-pounds oftorque.

Although the DC generator 10 and the gas turbine engine 5 are shown in aside-by-side relationship in the figure, it will be appreciated by thoseskilled in the art that an axially aligned (i.e., in-line)configuration, such as that shown in FIG. 7, may be preferred. Forexample, the DC generator 10 could be disposed longitudinally in frontof the gas turbine engine 5 such that primary shaft 15 extends forwardfrom the inlet 5A of engine 5 such that shaft 15 spins along an axiscommon to both the engine 5 and the DC generator 10. Such an in-lineconfiguration could make for a more efficient transfer of power from theengine 5 to the generator 10, as loss due to the right angled gearing isremoved. By positioning the generator 10 sufficiently ahead (forexample, ten to twelve feet) of the inlet 5A, the incidence of inletdistortion of the air entering the inlet 5A is reduced.

Depending on the size of the aircraft, single and multiple gas turbineengines may be utilized. Referring next to FIGS. 2A through 6B, top andelevation drawings of the incorporation of various embodiments of thepresent hybrid propulsion system 1 into numerous aircraft configurationsare shown. Aircraft employing the present hybrid propulsion system 1possess numerous attributes, including the ability to take offvertically and hover. In addition, helicopter configurations are capableof autogyro mode during flight, where in autogyro mode, directionalsteering can be accomplished by varying the speed of the jet engine 5and the rotors 30 in the outboard rotor configuration. Such could reduce(or outright do away with) the need for the tail rotor and its attendantweight and complexity.

Referring with particularity to FIGS. 2A and 2B, a single rotorhelicopter 100 is shown. Two gas turbine engines 5 are situated onlaterally opposing sides of the helicopter 100, and both (throughrespective shafts 15) can provide motive power to the bladed rotor 30.In addition, the transmission shafts 15 connect the engine 5 with the DCelectric generator 10. The bladed rotor 30 can be of any configurationdeemed appropriate for the aircraft; in the embodiment shown by way ofnon-limiting example, it may have from two to five blades to provideadequate lift. Landing rails 110 or wheels (not presently shown) may beused for ground engagement or transport. Another bladed rotor 70 issituated in the tail of helicopter 100 and is configured to providehorizontal stability.

Referring with particularity to FIGS. 3A and 3B, a larger helicopter 200with a front-to-back dual rotor configuration utilizing the hybridpropulsion system 1 discussed above is shown. Such a helicopter isparticularly well-suited to heavy lifting operations. The bladed rotors30 can be made to counter-rotate, thereby eliminating the need for ananti-torque vertical rotor such as the horizontally-stabilizing bladedrotor 70 of the device depicted in FIGS. 2A and 2B.

Referring with particularity to FIGS. 4A and 4B, a side-by-sideconfiguration of helicopter 300 is shown. Landing wheels 310 areincluded. In this configuration, a lateral frame 320 is included totransmit weight and lifting loads between the bladed rotors 30 and thefuselage of helicopter 300 Likewise, a side-by-side configuration ofhelicopter 600 is shown, where the propulsion system is configured asthe in-line variant of FIG. 7. Lateral frame 620 is used in a mannergenerally similar to lateral frame 320 of the device shown in FIGS. 4Aand 4B. Horizontal and vertical stability fins 625, similar to thoseused on fixed-wing aircraft, are also shown.

Referring with particularity to FIGS. 5A and 5B, a helicopter 400including a combination of the features of the embodiments of FIGS. 3A,3B, 4A and 4B are shown. Such a front-to-back quad rotor configurationutilizes eight of the hybrid propulsion systems.

Referring next to FIGS. 6A and 6B, as well as 8A and 8B, gyrocopters 500and 600 are shown. Gyrocopters share numerous features with helicopters(such as those shown in FIGS. 2A through 5B), but can not take offvertically nor hover the way a helicopter can. Nevertheless, theinfinite variability of the speed of the rotors and the turbine fordirectional steering in this invention makes a tail rotor unnecessary.Horizontal and vertical stability fins 525 (in the single rotor variantdepicted in FIGS. 6A and 6B) and 625 (in the twin rotor variant depictedin FIGS. 8A and 8B), such as those found on fixed-wing aircraft, areshown. Gyrocopter 600 highlights a particularly advantageousconfiguration that uses the in-line configuration of the propulsionsystem 1 that is depicted in FIG. 7.

While certain representative embodiments and details have been shown forpurposes of illustrating the invention, it will be apparent to thoseskilled in the art that various changes may be made without departingfrom the scope of the invention, which is defined in the appendedclaims.

What is claimed is:
 1. A method of providing motive power to anaircraft, said method comprising: coupling a hybrid propulsion system toa thrust producing device, said hybrid propulsion system comprising: agas turbine engine; an electric storage device configured to deliver anelectric current; an electrical generator responsive to said gas turbineengine such that power produced by the operation thereof is convertedinto electric current; and an electric motor selectively responsive tosaid electrical generator and said electric storage device; andoperating one of said gas turbine engine and said electric storagedevice such that power produced therefrom causes said electric motor tooperate said thrust producing device.
 2. The method of claim 1, whereinsaid thrust producing device operates by rotating in response to powerprovided to it by said electric motor.
 3. The method of claim 1, whereinsaid thrust producing device is configured to provide both lift andthrust to said aircraft.
 4. The method of claim 1, wherein said electricstorage device comprises at least one battery.
 5. The method of claim 4,wherein said operating one of said gas turbine engine and said at leastone battery comprises operating said at least one battery whileoperation of said gas turbine engine is substantially ceased.
 6. Themethod of claim 1, wherein said operating one of said gas turbine engineand said at least one battery comprises operating said gas turbineengine such that excess power produced therefrom is conveyed by itthrough said electrical generator to maintain said electric storagedevice in a substantially charged condition.
 7. The method of claim 1,wherein said aircraft is a helicopter or gyrocopter.
 8. The method ofclaim 1, wherein said operating one of said gas turbine engine and saidelectric storage device comprises providing computer-based automatedrobotic control to said hybrid propulsion system at least over a portionof a flight envelope of said aircraft.
 9. A method of reducing thein-flight thermal output of a rotary-winged aircraft propulsion system,said method comprising: coupling a hybrid propulsion system to a thrustproducing device, said hybrid propulsion system comprising: a gasturbine engine; an electric storage device configured to deliver anelectric current; an electrical generator responsive to said gas turbineengine such that power produced by the operation thereof is convertedinto electric current; and an electric motor selectively responsive tosaid electrical generator and said electric storage device; andoperating one of said gas turbine engine and said electric storagedevice such that power produced therefrom causes said electric motor tooperate said thrust producing device.
 10. An aircraft comprising: afuselage; a thrust-producing device coupled to said fuselage; and ahybrid propulsion system selectively coupled to said thrust-producingdevice such that cooperation therebetween provides motive power to saidaircraft, said hybrid propulsion system comprising: a gas turbine engineconfigured to operate over at least a first portion of said aircraft'sflight envelope; an electrical generator cooperative with said gasturbine engine such that mechanical power received therefrom by saidelectrical generator is converted to electric power; an electric storagedevice configured to operate over at least a second portion of saidaircraft's flight envelope; and an electric motor selectively coupled toat least one of said electrical generator and said electric storagedevice such that during said first portion of the aircraft's flightenvelope, said gas turbine engine and said thrust-producing deviceprovides at least a majority of said motive power, and during a secondportion of the aircraft's flight envelope, said electric storage deviceand said thrust-producing device provides at least a majority of saidmotive power.
 11. The aircraft of claim 10, wherein said electricstorage device comprises at least one battery.
 12. The aircraft of claim10, wherein said gas turbine engine comprises a shaft extendingtherefrom and a clutch cooperative with said shaft at least during saidfirst portion of the aircraft's flight envelope such that powergenerated in said gas turbine engine during such portion is transferredby said shaft to said electric generator.
 13. The aircraft of claim 10,wherein said electrical generator charges said electric storage deviceduring said first portion of the aircraft's flight envelope.
 14. Theaircraft of claim 10, wherein said electric motor is a direct currentelectric motor.
 15. The aircraft of claim 10, wherein saidthrust-producing device comprises a bladed rotor.
 16. The aircraft ofclaim 10, wherein during said second portion of the aircraft's flightenvelope, said gas turbine engine is substantially inoperable.
 17. Theaircraft of claim 10, further comprising a controller configured to varyoperation of said aircraft between said first and second portion of itsflight envelope.
 18. The aircraft of claim 17, wherein said controllercooperates with said hybrid propulsion system to provide computerizedcontrol over a substantial entirety of a flight envelope of saidaircraft.